Probe heater remaining useful life determination

ABSTRACT

A probe system includes a heater and a control circuit. The heater includes a resistive heating element routed through the probe. An operational voltage is provided to the resistive heating element to provide heating for the probe. The control circuit is configured to provide a test voltage different than the operational voltage and monitor a test current generated in the resistive heating element while providing the test voltage. The control circuit is further configured to detect micro fractures in the resistive heating element based on the test current.

BACKGROUND

The present invention relates generally to probes, and in particular to a system and method for determining a remaining useful life of aircraft sensor probes.

Probes are utilized to determine characteristics of an environment. In aircraft systems, for example, probes may be implemented on the external portions of the aircraft to aide in determination of conditions such as airspeed, Mach number and flight direction, among others. Due to the harsh conditions of flight, ice may build-up on portions of the probe. To combat this, heaters are implemented within the probe to prevent the formation of ice that may impact proper functionality of the probe.

When probes break down, they need to be replaced, often prior to a subsequent takeoff. The heating element of a probe is often the most life-limited part. Therefore, probes need to be replaced as soon as the heating element breaks down. It is desirable to predict a remaining useful life of the probe heating element in order to better predict maintenance needs of the probe itself.

SUMMARY

A system for an aircraft includes a probe and a control circuit. The probe includes a heater. The heater includes a resistive heating element routed through the probe. An operational voltage is provided to the resistive heating element to provide heating for the probe. The control circuit is configured to provide a low test voltage different from the operational voltage to the resistive heating element and monitor a test current through the resistive heating element while providing the low voltage. The control circuit is further configured to detect micro fractures in the resistive heating element based on the test current.

A method for detecting micro fractures in a resistive heating element of an aircraft probe includes providing, by a control circuit, a low test voltage to the resistive heating element of the aircraft probe; monitoring, by the control circuit, a test current while providing the low test voltage; detecting, by the control circuit, the micro fractures in the resistive heating element based on the test current; and determining a remaining useful life of the aircraft probe based upon detection of the micro fractures.

A probe system includes a heater and a control circuit. The heater includes a resistive heating element routed through the probe. An operational voltage is provided to the resistive heating element to provide heating for the probe. The control circuit is configured to provide a test voltage different than the operational voltage and monitor a test current generated in the resistive heating element while providing the test voltage. The control circuit is further configured to detect micro fractures in the resistive heating element based on the test current.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagram illustrating an aircraft that includes a plurality of probes.

FIG. 2 is a diagram of an aircraft probe that includes a heating element.

FIG. 3 is a diagram illustrating a heating element of an aircraft probe.

FIG. 4 is chart illustrating functions of a monitored characteristic of a plurality of probe heating elements over time.

FIGS. 5A and 5B are charts illustrating monitored current based on a low voltage for a heating element over time.

FIG. 6 is a chart illustrating a resonant frequency over time for a heating element of an aircraft probe.

FIGS. 7A and 7B are thermal images of an aircraft probe.

FIG. 8 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon detected micro-fractures.

FIG. 9 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon a monitored current draw over time by a heating element of a probe.

FIG. 10 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon a determined capacitance of a heating element of a probe.

FIG. 11 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon current leakage of a heating element of the probe.

FIG. 12 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon a resonant frequency of a heating element of the probe.

FIG. 13 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon thermal imaging of the probe.

FIG. 14 is a flowchart illustrating a method of determining a remaining useful life of a probe based upon an antenna response of a heater element of the probe.

DETAILED DESCRIPTION

A system and method for determining the remaining useful life of a probe is disclosed herein that includes monitoring characteristics of the probe over time. The probe, which may be an aircraft total-air-temperature (TAT) probe or any other probe, includes a resistive heating element, such as a heater wire, routed through the probe. Over time, as the heater wire ages, the heater wire may degrade, causing characteristics of the heater wire to change. These changing characteristics may be monitored and plotted over time, for example. The remaining useful life of the probe may then be determined and reported based upon the monitored characteristics.

FIG. 1 is a diagram illustrating aircraft 10 that includes a plurality of probes 12 a-12 n. While illustrated as a commercial aircraft, other vehicles, such as unmanned aerial vehicles, helicopters and ground vehicles may also include probes 12 a-12 n configured to sense characteristics of the environment. Probes 12 a-12 n may be any type of probe such as, but not limited to, pitot probes, TAT probes, angle-of-attack (AOA) probes and any other probes that may include a resistive heating element.

FIG. 2 illustrates aircraft probe 12 a that includes resistive heating element 14. While illustrated in FIG. 2 as a TAT probe 12 a, aircraft probe 12 a may be any other type of probe 12 a-12 n or sensing element. Probe 12 a is connected to receive control and power from control and interface circuit 16. Control and interface circuit 16 may be implemented local to probe 12 a (e.g., implemented as a “smart probe”) or remote of probe 12 a. Control and interface circuit 16 may include, for example, a microcontroller, programmable logic device, application integrated circuit (ASIC), or any other digital and/or analog circuitry.

Resistive heating element 14, which may be a heater wire, for example, may receive power directly, or through control and interface circuit 16, from aircraft power bus 18 to provide heating for probe 12 a. Power bus 18 may be any direct current (DC) or alternating current (AC) aircraft power bus. For example, resistive heating element 14 may receive power from a 28 Volt DC power bus. An operational current, based on the power received from power bus 18, flows through resistive heating element 14, which provides heating for probe 12 a. Control and interface circuit 16 may also be connected to aircraft avionics 20. Alternatively, control and interface circuit 16 may be implemented integral to aircraft avionics 20. Control and interface circuit 16 may be configured to provide data to, and receive data from, aircraft avionics 20.

Current sensors 22 a and 22 b may sense current flowing into, and out of, resistive heating element 14 at heating element input 26 a and heating element output 26 b, respectively. Current sensors 22 a and 22 b may provide a signal indicative of the sensed current at the respective locations to control and interface circuit 16. Temperature sensor 24 may be positioned to sense a temperature of probe 12 a and provide the sensed temperature to control and interface circuit 16. In other embodiments, a temperature may be estimated, for example, based upon sensed aircraft conditions and provided to control and interface circuit 16 from avionics 20. For example, avionics 20 may determine a present altitude and/or airspeed of aircraft 10 and may estimate the temperature of probe 12 a based upon, among other items, the present altitude and/or airspeed. Current sensors 22 a and 22 b may be any devices that sense current. In an embodiment, current sensors 22 a and 22 b may be non-contact current sensors such as, for example, a current transformer or Hall effect sensor.

Thermal imager 28 may be located integral to, or separate from, aircraft 10. Thermal imager 28 may be any device capable of receiving infrared radiation, for example, and providing an electrical output indicative of the received infrared radiation. While not illustrated as such, thermal imager 28 may be connected to communicate with control and interface circuit 16 and/or avionics 20 through a wired or wireless connection. This way, thermal images of probe 12 a may be obtained, analyzed and stored over time.

Radio-frequency (RF) antenna 30 is any antenna structure capable of emitting and/or receiving an RF signal. RF antenna 30 may be configured to receive power and emit RF radiation that may be received by heating element 14, for example. Heating element 14, which may include a large loop of heater wire having dielectric properties, may act as an antenna, capable of emitting and/or receiving RF energy. RF antenna 30 may be located integral to, or separate from, aircraft 10. RF antenna 30 may be connected to communicate with control and interface circuit 16 and/or avionics 20 through a wired or wireless connection (not shown). This way, control and interface circuit 16 may control RF antenna 30 to emit a plurality of RF signal frequencies, and/or may analyze a response of RF antenna 30 to an RF signal emitted by heater wire 14, for example.

Over time, resistive heating element 14 may degrade, and eventually break down such that current may no longer flow through resistive heating element 14 to provide heating for probe 12 a. Once resistive heating element 14 has broken down, aircraft probe 12 a must be repaired or replaced. A TAT probe, for example, may be utilized, among other things, to determine a Mach number for the aircraft. The Mach number may be needed for takeoff and thus, the TAT probe may be required to be functional prior to taking off. If the TAT probe is malfunctioning, it must be replaced, which may cause undesirable flight delays. If the remaining useful life of the TAT probe is known, the TAT probe can be replaced between flights or at another convenient time for repair, preventing delays or other costs incurred due to an unexpected failure of the probe.

FIG. 3 is a diagram illustrating an embodiment of heating element 14 of aircraft probe 12 a. Heating element 14, which is shown as a heater wire in the embodiment illustrated in FIG. 3, includes lead wire 40, heater element 42, insulation 44 and sheath casing 46. Sheath casing 46 may be a metallic sheath and thus, there may be a measurable capacitance between sheath casing 46 and lead wire 40. Ring oscillator 48 may be any oscillator circuit in which a capacitance may be utilized to drive an output frequency, for example. Capacitive measurement circuit 50 is any circuit capable of providing a value to control and interface circuit 16, for example, that allows control and interface circuit 16 to determine the capacitance between sheath casing 46 and lead wire 40. For example, capacitive measurement circuit 50 may be a resistor-capacitor (RC) or resistor-inductor-capacitor (RLC) circuit connected such that the capacitor for the RC or RLC circuit is the capacitance between lead wire 40 and metallic sheath 46. In some embodiments, ring oscillator 48 and capacitive measurement circuit 50 may be the same circuit.

During operation of probe 12 a, changes may occur, for example, due to hot and cold cycles as well as other varying environmental conditions such as temperature, pressure, humidity, environmental gases and aerosols, among others. These environmental conditions, in addition to the temperature cycling of probe 12 a, may cause a sealing of heating element 14 to become compromised, leading to foreign material leaking into insulation 44 and heater element 42. Heating element 14 may be oxidized and the dielectric material properties of heating element 14 may change. This may lead to certain characteristics of heating element 14 such as resistance and the capacitance between wire 40 and metallic sheath 46, among other characteristics, to change over time.

FIG. 4 illustrates functions 60 of a monitored characteristic over the life of several resistive heating elements 14 of several respective probes 12 a-12 n until breakdown 62. A heater wire, for example, may degrade as the heater wire ages. This degradation may cause changes in characteristics of heater wire 14 such as resistance and capacitance. For example, the current drawn through resistive heating element 14 may increase monotonically over the life of heating element 14. This increase is not linear, but rather follows exponential function 60, with a point 64 of interest. Point 64 may be the point on function 60, for example, at which the slope transitions from greater than 1 to less than 1. While the maximum normalized characteristic and life of each resistive heating element 14 at breakdown 62 may fluctuate, exponential function 60 of the heater characteristic over the life of each resistive heating element 14 remains similar.

Functions 60 shown in FIG. 4 may be obtained, for example, through testing of probes 12 a-12 n. For a selected characteristic of heating element 14, such as current draw, capacitance, leakage current, thermal response, or other characteristic, function 60 may be determined through the testing of probes 12 a-12 n. Current may be cycled, for example, until heating element 14 of a respective test probe 12 a-12 n breaks down. A characteristic of heating element 14 may be monitored each test cycle and plotted over time until breakdown 62. The plots of the characteristic may then be utilized to determine function 60. Function 60 may then be utilized during normal operation of a probe 12 a-12 n to determine, for example, a half-life estimate of heating element 14. In other embodiments, other algorithms may be utilized to determine the remaining useful life of heating element 14 based upon the plotted data.

In an example embodiment, current may be sampled using current sensor 22 a or 22 b at any time during flight and provided to control and interface circuit 16 several times each flight. This sampled current may be stored in a memory of control and interface circuit 16, for example, or some other memory or storage device. Because current is directly affected by temperature, a temperature may also be sensed using temperature sensor 24 and stored along with a respective sensed current so that the current may be normalized with respect to temperature. Alternatively, a temperature may be estimated using data from avionics 20, for example. Control and interface circuit 16 may utilize the sensed or estimated temperature to directly normalize the current prior to storage, or may store the two values for later normalization.

Function 60, determined during testing, may be used in conjunction with the stored normalized current to determine the half-life estimate of heating element 14. For example, the stored normalized current may be plotted by control and interface circuit 16 such that a present slope of the plot may be determined. This present slope may be utilized, for example, to detect point 64 by detecting that the slope of the plot has transitioned from greater than 1 to less than 1, or any other point on function 60. Once point 64 is detected, for example, the half-life of heating element 14 may be determined.

FIGS. 5A and 5B are charts illustrating monitored current based upon a normal operating voltage and a low voltage for heating element 14 over time. FIG. 5A illustrates both a current draw 70 for an operating voltage and current draw 72 for a low voltage. FIG. 5B shows a zoomed in view of current draw 72 for a low voltage. In an embodiment, the operating voltage may be, for example, 28 Volts while the low voltage may be, for example, 0.1 Volts.

At the end of the life of heating element 14, thermal fatigue causes heater wire 40 to fracture and gradually become electrically open, which increases the resistance of heating element 14. At low voltage, electron tunneling is the main mechanism for electrical current conduction. Arcing that occurs at 0.1 Volts may indicate a micro fracture with a gap on the order of 1 nanometer (nm), while arcing that occurs at 28 Volts may correspond to an approximately 0.3 micrometer (um) gap. Thus, micro fractures may be detected by monitoring a current response using a low voltage prior to failure of heating element 14.

As seen in FIG. 5A, the current drawn during normal operation (e.g., receiving 28 Volts from power bus 18) remains substantially greater than zero over all cycles. As seen in FIG. 5B, the current drawn with 0.1 Volts begins to degrade to, and remain at, zero while the current drawn during normal operation continues to be substantially greater than zero. This degradation of current to zero at low voltage indicates that a micro fracture is present in heating element 14. This micro fracture may eventually grow in size, ultimately causing failure of probe 12 a. By detecting these micro fractures early, failure of probe 12 a may be predicted such that a remaining useful life of heating element 14, and in turn, probe 12 a, may be determined.

The amount of remaining useful life of probe 12 a following detection of a micro fracture may be determined, for example, through testing of probes 12 a-12 n. Current may be cycled, for example, until heating element 14 of a respective test probe 12 a-12 n breaks down. Between test cycles, a low voltage may be provided to test probes 12 a-12 n. While the low voltage is supplied, a current may be sensed and stored. This way, the cycle at which a micro fracture on the order of 1 nm is detected may be determined. Following detection of the micro fracture, once the respective test probe 12 a-12 n breaks down, a percentage of life after detection of the micro fracture may be determined.

During normal operation of probes 12 a-12 n, a low voltage may be provided, for example, by control and interface circuit 16 between flights of aircraft 10, or at any other time during which heating element 14 is not receiving an operational voltage. Current sensor 22 a or 22 b may be utilized by control and interface circuit 16 to obtain a sensed current while providing the low voltage. Upon detection of the sensed current going to zero at low voltage, a remaining useful life may be determined based upon the percentage of life determined during testing of the probes 12 a-12 n.

FIG. 6 is a chart illustrating a resonant frequency over time for heating element 14 of aircraft probe 12 a. FIG. 6 also illustrates a curve 80 fit to the resonant frequency plot over time. As seen in FIG. 3, a capacitance exists between metallic sheath 46 and lead wire 40. While in operation, changes to heating element 14 happen due to hot and cold cycles in addition to varying environmental conditions such as temperature, pressure, humidity, environmental gases, and aerosols, among others. These conditions, in addition to the temperature cycling of probe 12 a, cause the probe heater sealing to be compromised, which may lead to various materials leaking into insulation 44 and heater element 42. Wire 40 may be oxidized and dielectric material properties may change, leading to a change in capacitance and resistance of heating element 14. Because the capacitance changes, the resonant frequency also changes.

In an embodiment, to obtain the resonant frequency, the capacitance between metallic sheath 46 and lead wire 40 may be swept with frequencies ranging from 1 kilohertz (KHz) to 100 megahertz (MHz) by control and interface circuit 16, for example, through capacitive measurement circuit 50. The peak frequency response and bandwidth may be identified by control and interface circuit 16. This peak frequency response may be monitored and stored over time and plotted as shown in FIG. 6. In one embodiment, a function 80 obtained during testing of probes 12 a-12 n, for example, may then be utilized in conjunction with the stored and plotted resonant frequency to determine a remaining useful life of probe 12 a. For example, it may be determined when the slope of the plotted resonance transitions from greater than −1 to less than −1, or when the slope transitions from a negative value to a positive value, as seen in FIG. 6. Determination of a present point on function 80 allows control and interface circuit 16 to determine a remaining useful life of probe 12 a.

In another embodiment, algorithms that separate data indicative of a healthy probe and data indicative of an increasingly unhealthy probe may be executed utilizing various signal processing techniques to determine the remaining useful life of heating element 14. For example, time-frequency analysis, machine learning, index theory and other signal processing techniques may be utilized to determine remaining useful life of heating element 14 based upon the monitored resonant frequency.

In another embodiment, ring oscillator 48 may be utilized with probe 12 a as the driving element of the output frequency of ring oscillator 48. Ring oscillator 48, or other oscillator circuit, has an output frequency dependent upon the structure of the circuit and the input voltage to the circuit. By controlling the input voltage, the output frequency of ring oscillator 48 can be made dependent solely upon the capacitance between metallic sheath 46 and lead wire 40. As this capacitance changes as heating element 14 degrades, so does the output frequency of ring oscillator 48. The changing output frequency may be stored and plotted over time to determine a remaining useful life of probe 12 a. For example, testing of probes 12 a-12 n may result in a function of the output frequency similar to that shown in FIG. 4 or 6. This function, in conjunction with the stored and plotted output frequency, may be utilized to determine the remaining useful life of probe 12 a during normal operation.

In another embodiment, RF antenna 30 may be controlled to sweep a wide range of frequencies. For example, control and interface circuit 16 may provide AC power to antenna 30 at varying frequencies to facilitate emission of RF radiation at a plurality of frequencies from antenna 30. The S12 parameter for heating element 14, which is a measure of the power received at heating element 14 from the RF emission of antenna 30, may be determined and monitored by control and interface circuit 16.

This antenna response of heating element 14 may be analyzed by control and interface circuit 16 to determine, for example, a resonant frequency response of heating element 14. The antenna properties of heating element 14 may be dependent upon, among other things, length and shape of heating element 14, and dielectric properties of heating element 14. These dielectric properties may include, for example, permittivity, permeability, homogeneity and thickness, among others. As the wire ages and degrades, the dielectric properties of heating element 14 may change, which may lead to a change in the resonant frequency of heating element 14.

The determined resonant frequency may be monitored and plotted over time by control and interface circuit 16. Testing of probes 12 a-12 n may be utilized to determine a function at which the resonant frequency changes over time. This function may substantially follow those shown in FIG. 4 or 6, for example. This function may then be utilized to determine, for example, a half-life of heating element 14 base upon the determined resonant frequency during normal operation of probe 12 a. In other embodiments, an algorithm may be executed using signal processing techniques to determine the remaining useful life of probe 12 a based upon the resonant frequency changes over time.

In another embodiment, control and interface circuit 16 may provide AC power to heating element 14 at varying frequencies to facilitate emission of RF radiation at a plurality of frequencies from heating element 14. Control and interface circuit 16 may then monitor the S12 parameter of RF antenna 30 to determine a resonant frequency, which may be monitored and analyzed over time to determine a half-life of heating element 14.

FIGS. 7A and 7B are representatives of thermal images of aircraft probe 12 a. These images may be obtained, for example, using thermal imager 28 (FIG. 2). FIG. 7A illustrates probe 12 a at a beginning of its lifetime, while FIG. 7B illustrates probe 12 a at a later point in its lifetime. The thermal images may be utilized, for example, to determine a thermal response of heating element 14 over time. Points of interest 90 a and 92 a show a thermal response of heating element 14 at the beginning of life of heating element 14, while points of interest 90 b and 92 b show a thermal response of heating element 14 at a later point in life. While illustrated as images captured at various view angles with respect to probe 12 a, any number of thermal images at any angle with respect to probe 12 a may be obtained by imager 28. Other points of interest for heating element 14 may also be monitored separately, or in conjunction with, points 90 a, 90 b, 92 a and 92 b to determine the thermal response of heating element 14.

As seen in FIG. 7B, the temperature at points 90 b and 92 b of heating element 14 has increased from the temperature at points 90 a and 92 a shown in FIG. 7A. This increase may be due to changes in heating element caused by degradation of heating element 14 over time. The increase in temperature over time may substantially follow an exponential function similar to those illustrated in FIG. 4. Because of this, the thermal response of heating element 14 (e.g., the increase in temperature seen at points 90 b and 92 b) may be stored and plotted over time during normal operation of probe 12 a. An exponential function determined during testing of probes 12 a-12 n, for example, may then be used in conjunction with the stored and plotted thermal response to determine a half-life of heating element 14.

FIG. 8 is a flowchart illustrating method 100 of determining a remaining useful life of probe 12 a based upon detected micro-fractures. At step 102, a low voltage is provided to heating element 14 of probe 12 a. This low voltage may be provided at any time, such as right after power down of probe 12 a. The low voltage may be approximately 0.1 V, which may be low enough to detect micro fractures on the order of 1 nm.

A step 104, while the low voltage is being supplied to heating element 14, current is sensed by one of current sensors 22 a and 22 b. The sensed current may be provided to, and stored by, control and interface circuit 16. The process in steps 102 and 104 is repeated over time, and the sensed current is monitored by control and interface circuit 16 at step 106. At step 108, micro fractures in heating element 14 are detected based upon the monitored current. For example, if the monitored current has dropped to zero, as illustrated in FIG. 5B, a micro fracture is detected by control and interface circuit 16.

Following detection of an approximately 1 nm micro fracture, the remaining useful life of heating element 14 may be estimated. This estimation may be based upon testing of heating elements 14, for example. Probes 12 a-12 n may be tested to determine an average remaining useful life of heating element 14 following detection of a micro fracture of approximately 1 nm.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 110. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

FIG. 9 is a flowchart illustrating method 120 of determining a remaining useful life of a probe 12 a-12 n based upon monitored current drawn over time by a respective heating element 14. At step 122, an exponential function for a respective probe 12 a-12 n may be determined, for example, through testing of similar probes 12 a-12 n. The test probes may have a test current provided to a respective resistive heating element 14 during a plurality of test cycles. The current may be sensed at common points during each of the plurality of test cycles and the sensed currents may then be plotted and functions 60 may be determined as shown in FIG. 4 for the respective resistive heating elements 14. This exponential function may then be used for all similar probes 12 a-12 n during normal probe operation in order to determine a half-life of a respective resistive heating element 14, for example.

During normal operation of probes 12 a-12 n, current through resistive heating element 14 may fluctuate based upon the operating point of probe 12 a-12 n. Therefore, it may be desirable at step 124 to sense the operational current at a similar point of operation of heating element 14. For example, upon initial power-on of resistive heating element 14, the current may rise to a peak current. As resistive heating element 14 increases in temperature, the current through resistive heating element 14 will decrease to a lower, steady-state current. Thus, current may be sampled and stored consistently at the peak value, at the steady-state value, or at some other expected value, for example.

Temperature may also be sensed and provided to control and interface circuit 16 using temperature sensor 24. At step 126, control and interface circuit 16 may normalize the sensed current using the sensed temperature. Steps 124 and 126 may be repeated and the normalized sensed current may be plotted over several flights, for example, to establish a curve fit. The curve fit may substantially follow the exponential function determined, for example, at step 122. At step 128, the present slope of the sensed current may be determined based upon the exponential function. This slope may then be utilized to determine the half-life of resistive heating element 14. By knowing the half-life of resistive heating element 14, the remaining useful life of resistive heating element 14 may be determined by control and interface circuit 16, for example.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 130. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

FIG. 10 is a flowchart illustrating method 140 of determining a remaining useful life of probes 12 a-12 n based upon a determined capacitance of a respective heating element 14. At step 142, an exponential function for a respective probe 12 a-12 n may be determined, for example, through testing of similar test probes. The test probes may have a test current provided to a respective resistive heating element 14 during a plurality of test cycles to simulate the life of the respective probe 12 a-12 n. The current may be cycled until failure of each respective probe 12 a-12 n. The capacitance of each respective heating element 14 may be determined during each cycle, for example. An exponential function, such as those shown in FIG. 4, may be determined based upon the determined capacitances. This exponential function may then be used for all similar probes 12 a-12 n during normal probe operation in order to determine a half-life of a respective resistive heating element 14, for example.

At step 144, during normal operation of probes 12 a-12 n, capacitive measurement circuit 50 may be utilized to determine a capacitance between metallic sheath 46 and lead wire 40. The capacitance and capacitance measurement circuit 50 may be directly affected by temperature. At step 146, temperature may be sensed and provided to control and interface circuit 16 using temperature sensor 24. Steps 144 and 146 may be repeated throughout the life of probe 12 a and control and interface circuit 16 may normalize the determined capacitance using the sensed temperature and may plot the normalized determined capacitance over several flights, for example, to establish a curve fit. The curve fit may substantially follow the exponential function determined, for example, at step 142. At step 148, the present slope of the determined capacitance may be determined based upon the exponential function. This slope may then be utilized to determine the half-life of resistive heating element 14. By knowing the half-life of resistive heating element 14, the remaining useful life of resistive heating element 14 may be determined by control and interface circuit 16, for example.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 150. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

FIG. 11 is a flowchart illustrating method 160 of determining a remaining useful life of probes 12 a-12 n based upon a measured leakage current for a respective heating element 14. At step 162, an exponential function for a respective probe 12 a-12 n may be determined, for example, through testing of similar probes 12 a-12 n. The test probes may have a test current provided to a respective resistive heating element 14 during a plurality of test cycles to simulate the life of the respective probe 12 a-12 n. Current may be sampled by both current sensors 22 a and 22 b and provided to control and interface circuit 16. The current from current sensor 22 b for example, may be subtracted from the current from current sensor 22 a to determine a leakage current of resistive heating element 14. The current may be cycled until failure of each respective probe 12 a-12 n at which time the exponential function, similar to those illustrated in FIG. 4, may be determined.

At step 164, during normal operation of probes 12 a-12 n, leakage current from resistive heating element 14 may be determined based upon sensed current from both current sensors 22 a and 22 b. Current through resistive heating element 14 may fluctuate based upon the operating point of probe 12 a-12 n. Therefore, it may be desirable to determine the leakage current at a similar point of operation of heating element 14 each time the current is sensed and stored. For example, upon initial power-on of resistive heating element 14, the current may rise to a peak current. As resistive heating element 14 increases in temperature, the current through resistive heating element 14 will decrease to a lower, steady-state current. Thus, current may be sampled and stored consistently at the peak value, at the steady-state value, or at some other expected value, for example.

Temperature may also be sensed and provided to control and interface circuit 16 using temperature sensor 24. At step 166, control and interface circuit 16 may normalize the determined leakage current using the sensed temperature. Steps 164 and 166 may be repeated and control and interface circuit 16 may plot the normalized leakage current over several flights, for example, to establish a curve fit. The curve fit may substantially follow the exponential function determined, for example, at step 162. At step 168, the present slope of the determined leakage current may be determined based upon the exponential function. This slope may then be utilized to determine the half-life of resistive heating element 14. By knowing the half-life of resistive heating element 14, the remaining useful life of resistive heating element 14 may be determined by control and interface circuit 16, for example.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 170. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

FIG. 12 is a flowchart illustrating method 180 of determining a remaining useful life of probes 12 a-12 n based upon a resonant frequency of a respective heating element 14. At step 182, an exponential function for a respective probe 12 a-12 n may be determined, for example, through testing of similar test probes. The exponential function may be similar to that shown in FIG. 6. The test probes may have a test current provided to a respective resistive heating element 14 during a plurality of test cycles to simulate the life of the respective probe 12 a-12 n. The current may be cycled until failure of each respective probe 12 a-12 n. The resonant frequency based on the capacitance of each respective heating element 14 may be determined during each cycle, for example, using capacitive measurement circuit 50. Alternatively, an output frequency of oscillator circuit 48 driven by the capacitance of heating element 14 may be determined during each cycle. An exponential function, such as that shown in FIG. 4 or FIG. 6, may be determined based upon the determined resonant or output frequency. This exponential function may then be used for all similar probes 12 a-12 n during normal probe operation in order to determine a half-life of a respective resistive heating element 14, for example.

At step 184, during normal operation of probes 12 a-12 n, capacitive measurement circuit 50 may be used to determine a resonant frequency of heating element 14 or ring oscillator circuit 48 may be utilized to determine an output frequency of ring oscillator 48 driven by the capacitance of heating element 14. At step 186, temperature may also be sensed and provided to control and interface circuit 16 using temperature sensor 24. Control and interface circuit 16 may normalize the determined resonant frequency using the sensed temperature. Steps 184 and 186 may be repeated and control and interface circuit 16 may plot the normalized resonant or output frequency over several flights, for example, to establish a curve fit. While theoretical capacitance is not directly affected by temperature, in practice both the capacitance and ring oscillator circuit 48 may be directly affected by temperature. The curve fit may substantially follow the exponential function determined, for example, at step 182. At step 188, the present slope of the determined resonance may be determined based upon the exponential function. This slope may then be utilized to determine the half-life of resistive heating element 14. By knowing the half-life of resistive heating element 14, the remaining useful life of resistive heating element 14 may be determined by control and interface circuit 16, for example.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 190. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

FIG. 13 is a flowchart illustrating method 200 of determining a remaining useful life of probes 12 a-12 n based upon thermal imaging of a respective heating element 14. At step 202, an exponential function for a respective probe 12 a-12 n may be determined, for example, through testing of similar test probes. The exponential function may be similar to that shown in FIG. 4. The test probes may have a test current provided to a respective resistive heating element 14 during a plurality of test cycles to simulate the life of the respective probe 12 a-12 n. The current may be cycled until failure of each respective probe 12 a-12 n. Thermal images may be taken by thermal imager 28 for each respective heating element 14 each cycle, for example. An exponential function, such as that shown in FIG. 4, may be determined based upon the thermal images. This exponential function may then be used for all similar probes 12 a-12 n during normal probe operation in order to determine a half-life of a respective resistive heating element 14, for example.

At step 204, during normal operation of probes 12 a-12 n, thermal imager 28 may be utilized to obtain thermal images of heating element 14. At step 206, temperature may also be sensed and provided to control and interface circuit 16 using temperature sensor 24. Control and interface circuit 16 may normalize the thermal data using the sensed temperature. Control and interface circuit 16 may also normalize the thermal image data to compensate for probe shape and surface emissivity, for example. Steps 204 and 206 may be repeated and control and interface circuit 16 may plot the normalized thermal data over several flights, for example, to establish a curve fit. The curve fit may substantially follow the exponential function determined, for example, at step 202. At step 208, the present slope of the determined thermal data be determined based upon the exponential function. This slope may then be utilized to determine the half-life of resistive heating element 14. By knowing the half-life of resistive heating element 14, the remaining useful life of resistive heating element 14 may be determined by control and interface circuit 16, for example.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 210. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

FIG. 14 is a flowchart illustrating method 220 of determining a remaining useful life of probes 12 a-12 n based upon an antenna response of a respective heating element 14. At step 222, an exponential function for a respective probe 12 a-12 n may be determined, for example, through testing of similar test probes. The exponential function may be similar to that shown in FIG. 4. The test probes may have a test current provided to a respective resistive heating element 14 during a plurality of test cycles to simulate the life of the respective probe 12 a-12 n. The current may be cycled until failure of each respective probe 12 a-12 n. RF antenna 30 may be utilized during each cycle, for example, to sweep frequencies of RF radiation to heating element 14. An exponential function, such as that shown in FIG. 4, may be determined based upon a detected resonant frequency of heater element 14 in response to the RF radiation from antenna 30 over each cycle. This exponential function may then be used for all similar probes 12 a-12 n during normal probe operation in order to determine a half-life of a respective resistive heating element 14, for example.

At step 224, during normal operation of probes 12 a-12 n, RF antenna 30 may be utilized to provide RF radiation to heating element 14. At step 226, control and interface circuit 16 may monitor the S12 parameter of heating element 14 to determine a resonant frequency response of heating element 14 to the RF radiation. In an alternative embodiment, heating element 14 may be energized to provide RF radiation to antenna 30 and control and interface circuit 16 may monitor the S12 parameter of antenna 30 to determine a resonant frequency response of antenna 30.

Steps 224 and 226 may be repeated and control and interface circuit 16 may plot the resonant frequency response over several flights, for example, to establish a curve fit. The curve fit may substantially follow the exponential function determined, for example, at step 222. At step 228, the present slope of the determined RF response may be determined based upon the exponential function. This slope may then be utilized to determine the half-life of resistive heating element 14. By knowing the half-life of resistive heating element 14, the remaining useful life of resistive heating element 14 may be determined by control and interface circuit 16, for example. During steps 224 and 226, the RF response may be normalized to account for, among other things, temperature. This may be accomplished using data from temperature sensor 24, or an estimate temperature based upon data from avionics 20, for example.

Once the remaining useful life of resistive heating element 14 is determined by control and interface circuit 16, the remaining useful life may be reported at step 230. This report may be made to the cockpit through avionics 20, or to some other computer system. By reporting the remaining useful life, probes 12 a-12 n may be replaced prior to breaking down and thus, unnecessary flight delays may be avoided.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments of the present invention.

A system for an aircraft includes a probe and a control circuit. The probe includes a heater. The heater includes a resistive heating element routed through the probe. An operational voltage is provided to the resistive heating element to provide heating for the probe. The control circuit is configured to provide a low test voltage different from the operational voltage to the resistive heating element and monitor a test current through the resistive heating element while providing the low voltage. The control circuit is further configured to detect micro fractures in the resistive heating element based on the test current.

The system of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:

A further embodiment of the foregoing system, further comprising a current sensor configured to sense a sensed current through the resistive heating element and provide the sensed current to the control circuit.

A further embodiment of any of the foregoing systems, wherein the control circuit is configured to detect micro fractures while providing the low test voltage if the sensed current is zero.

A further embodiment of any of the foregoing systems, wherein the low test voltage is 0.1 Volts.

A further embodiment of any of the foregoing systems, wherein the operational voltage is received from a 28 Volt direct current power bus.

A further embodiment of any of the foregoing systems, wherein the probe is a total air temperature probe configured to sense a total air temperature outside the aircraft.

A further embodiment of any of the foregoing systems, wherein the control circuit is configured to monitor the test current over a plurality of flights of the aircraft and provide an indication upon detection of the micro fractures in the resistive heating element.

A further embodiment of any of the foregoing systems, wherein the control circuit is configured to determine a remaining useful life of the resistive heating element based upon detection of the micro fractures.

A method for detecting micro fractures in a resistive heating element of an aircraft probe includes providing, by a control circuit, a low test voltage to the resistive heating element of the aircraft probe; monitoring, by the control circuit, a test current while providing the low test voltage; detecting, by the control circuit, the micro fractures in the resistive heating element based on the test current; and determining a remaining useful life of the aircraft probe based upon detection of the micro fractures.

The method of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:

A further embodiment of the foregoing method, wherein the aircraft probe is a total air temperature probe configured to sense a total air temperature outside the aircraft.

A further embodiment of any of the foregoing methods, wherein monitoring, by the control circuit, the test current while providing the low test voltage includes sensing, by a current sensor, a sensed test current of the resistive heating element; and providing the sensed test current to the control circuit.

A further embodiment of any of the foregoing methods, by the control circuit, the low test voltage to the resistive heating element of the aircraft probe includes providing 0.1 Volts to the resistive heating element.

A further embodiment of any of the foregoing methods, wherein detecting, by the control circuit, the micro fractures in the resistive heating element based on the test current includes detecting, by the control circuit, that the sensed test current is zero.

A probe system includes a heater and a control circuit. The heater includes a resistive heating element routed through the probe. An operational voltage is provided to the resistive heating element to provide heating for the probe. The control circuit is configured to provide a test voltage different than the operational voltage and monitor a test current generated in the resistive heating element while providing the test voltage. The control circuit is further configured to detect micro fractures in the resistive heating element based on the test current.

The probe system of the preceding paragraph can optionally include, additionally and/or alternatively, any one or more of the following features, configurations and/or additional components:

A further embodiment of the foregoing probe system, further comprising a current sensor configured to sense a sensed current through the resistive heating element and provide the sensed current to the control circuit.

A further embodiment of any of the foregoing probe systems, wherein the control circuit is configured to detect micro fractures while providing the low test voltage if the sensed current is zero.

A further embodiment of any of the foregoing probe systems, wherein the low test voltage is 0.1 Volts.

A further embodiment of any of the foregoing probe systems, wherein the operational voltage is received from a 28 Volt direct current power bus.

A further embodiment of any of the foregoing probe systems, wherein the probe is a total air temperature probe configured to sense a total air temperature outside the aircraft.

A further embodiment of any of the foregoing probe systems, wherein the control circuit is configured to monitor the test current over a plurality of flights of the aircraft and provide an indication upon detection of the micro fractures in the resistive heating element.

While the invention has been described with reference to an exemplary embodiment(s), it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the invention. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the invention without departing from the essential scope thereof. Therefore, it is intended that the invention not be limited to the particular embodiment(s) disclosed, but that the invention will include all embodiments falling within the scope of the appended claims. 

The invention claimed is:
 1. A system for an aircraft, the system comprising: a probe that includes a heater, wherein the heater includes a resistive heating element routed through the probe, and wherein an operational voltage is provided to the resistive heating element to provide heating for the probe; and a control circuit configured to provide a low test voltage different from the operational voltage to the resistive heating element, wherein the control circuit is configured to monitor a test current through the resistive heating element while providing the low voltage; wherein the control circuit is further configured to detect micro fractures in the resistive heating element based on the test current and determine a remaining useful life of the resistive heating element based upon detection of the micro fractures in the resistive heating element during a plurality of test cycles and functions determined during the plurality of test cycles.
 2. The system of claim 1, further comprising a current sensor configured to sense a sensed current through the resistive heating element and provide the sensed current to the control circuit.
 3. The system of claim 2, wherein the control circuit is configured to detect micro fractures while providing the low test voltage if the sensed current is zero.
 4. The system of claim 1, wherein the low test voltage is 0.1 Volts.
 5. The system of claim 4, wherein the operational voltage is received from a 28 Volt direct current power bus.
 6. The system of claim 1, wherein the probe is a total air temperature probe configured to sense a total air temperature outside the aircraft.
 7. The system of claim 1, wherein the control circuit is configured to monitor the test current over a plurality of flights of the aircraft and provide an indication upon detection of the micro fractures in the resistive heating element.
 8. The system of claim 7, wherein the control circuit is configured to determine a remaining useful life of the resistive heating element based upon detection of the micro fractures.
 9. A method for detecting micro fractures in a resistive heating element of an aircraft probe, the method comprising: providing, by a control circuit, a low test voltage to the resistive heating element of the aircraft probe; monitoring, by the control circuit, a test current while providing the low test voltage; detecting, by the control circuit, the micro fractures in the resistive heating element based on the test current; and determining a remaining useful life of the aircraft probe based upon detection of the micro fractures in the resistive heating element during a plurality of test cycles and functions determined during the plurality of test cycles.
 10. The method of claim 9, wherein the aircraft probe is a total air temperature probe configured to sense a total air temperature outside the aircraft.
 11. The method of claim 9, wherein monitoring, by the control circuit, the test current while providing the low test voltage comprises: sensing, by a current sensor, a sensed test current of the resistive heating element; and providing the sensed test current to the control circuit.
 12. The method of claim 11, wherein providing, by the control circuit, the low test voltage to the resistive heating element of the aircraft probe comprises providing 0.1 Volts to the resistive heating element.
 13. The method of claim 12, wherein detecting, by the control circuit, the micro fractures in the resistive heating element based on the test current comprises detecting, by the control circuit, that the sensed test current is zero.
 14. A probe system comprising: a heater comprising a resistive heating element routed through the probe, wherein an operational voltage is provided to the resistive heating element to provide heating for the probe; and a control circuit configured to provide a test voltage different than the operational voltage and monitor a test current generated in the resistive heating element while providing the test voltage; wherein the control circuit is further configured to detect micro fractures in the resistive heating element based on the test current and determine a remaining useful life of the resistive heating element based upon detection of micro fractures in the resistive heating element during a plurality of test cycles and functions during the plurality of test cycles.
 15. The probe system of claim 14, further comprising a current sensor configured to sense a sensed current through the resistive heating element and provide the sensed current to the control circuit.
 16. The probe system of claim 15, wherein the control circuit is configured to detect micro fractures while providing the low test voltage if the sensed current is zero.
 17. The probe system of claim 14, wherein the low test voltage is 0.1 Volts.
 18. The probe system of claim 17, wherein the operational voltage is received from a 28 Volt direct current power bus.
 19. The probe system of claim 14, wherein the probe is a total air temperature probe configured to sense a total air temperature outside the aircraft.
 20. The probe system of claim 14, wherein the control circuit is configured to monitor the test current over a plurality of flights of the aircraft and provide an indication upon detection of the micro fractures in the resistive heating element. 